Rotary bladed fluid flow machines



Dec. 30, 1958 A. D. s. CARTER 2,866,619

ROTARY BLADED FLUID FLOW MACHINES Filed March 5, 1952 2 Sheets-Sheet l aInventor Dec. 30, 1958 A. D. s. CARTER ROTARY BLADED FLUID FLOW MACHINES2 Sheets-Sheet 2 Filed March 5, 1952 i R 9k 2 Invenfas 2 United StatesPatent ROTARY BLADED FLUID FLOW MACHINES Alfred Denis Snowdon Carter,Farnb'orough, England, assignor to Power Jets (Research and Development)Limited, London, England, a Britishcompany Application March 5, 1952,Serial No. 274,950

Claims priority, application Great Britain'March 9, 1951 3 Claims. (Cl.253-65) This invention relates to rotary bladed -fluid flow machines,more particularly to turbines, and has for an object the improvement ofthe efliciency'thereof by reducing energy losses associated with thepassage of ffluid through the machine.

In general, the energy losses associated with a row of cambered bladessuch as are used in compressors and turbines are attributable to twomain causes, namely frictional drag between the fluid and surfaces 'of"the blades and bounding walls, known as primary losses, and dissipationof kinetic energy of the fluid invarious'turbulent flows inducedtherein, known as secondary. losses, and it is with the reduction of thelatter thatthe present invention is concerned.

The invention proposes broadly that the row or last row of blading in amachine of the kind under consideration intended primarily for energyconversion (i. e. the working blades) be immediately succeeded, in thedirection of fluid flow, by a row of blades with respect to which thisworking blade'row is rotatable having loss characteristics which are lowin relation to those of'the next preceding row of blades, i. e. asexplained in'more detail hereinafter, blades which are at most of verylow cambered section compared with the blades of the preceding row.

The invention arises from a consideration of the probable relativemagnitudes of primary and secondary losses in various cases and thefactors influencing their magnitude as inferred from knowledge atpresent available, this inferential assessment being obligatory in theabsence of any technique for direct but separate measurement of the twoforms of loss. I

An indication of the magnitude of these'losses .is afforded by the factthat a typical stage of compressor blading comprising an intermediatestage of a'multi-stage machine would have an efficiency, at best, ofapproximately 90 percent. Practical experiencerhas led to' thecomputation of primary and secondary losses as accounting for 6 percentand 4 percent respectively of the 10 percent total loss of efliciency,and in order to' afford a basis for the accurate estimation of secondarylosses in such a typical compressor stage the secondary losses have beenconveniently expressed as giving rise to a'co-e'fiicient of drag inducedthereby, C and the following empirical formula propounded:

where C is the co-efficient of lift of the blades of'the compressorstage.

The turbulence giving rise to the secondary losses has been found intests on cascades of blades under laboratory conditions to beprincipally in the'form of vortices extending downstream from thetrailing edge-of the-blades, a'pairof oppositely rotating vortices beingsituated one toward each end of each blade adjacent the bounding walls.Analysis of the circumstances surrounding the formation of this vortexsystem has led t'o'anacc'eptable explanation and understanding of thephenomena-an'd ice to a theoretical basis for the-assessment of. thesecondary losses, again expressedffor convenience as a co-eflicient ofdrag. induced thereby, .by identifying the phenomena with a vortex'systemextending from the blade an infinite distance downstream. Thusthe-followingexpression has been derived wheres/c is the pitch/chordratio of the blades in .the cascade,.and h'/h is the ratio of the spanof the axes of the pair of vortices extending from each blade to thespan of the blades. Typical values for these ratios in a compressorstage would 'be of the order 0.88 and 0.75 respectively so thatingeneral by substitution in the. above equation;

Cm=0.0532 CLZ Thus both practical and theoretical assessments of 'thesecondary loss induced drag co-eflicient are dependent on C the liftco-efiicient; this is to be expected since the turbulence of the vortex'system'is,-in general, intensified with increasing blade camber, whichalso gives rise to increased lift co-efficients. In respect of actualmagnitude, however, the practical and theoretical estimates do notagree, the latter being approximately 3 times the former, and there isconsequently a natural tendency to doubt the theoretical estimate.

- On the other hand, tests have beencarried out, for the purpose ofmeasuring the variation of total losses in a single stage turbine withvarying blade tip clearances, from which some indirect support for-thevalidity of the theoretical estimate of secondary loss induced drag maybe deduced. Since a variation in tip clearance'would seemingly notaffect appreciably the frictional losses of blades and bounding walls,the variation in total losses would be expected to give a directindication of the variation of secondary losses due to re-adjustment ofthe span of the vortex system upon variation of the tip clearance. Onthis basis, it is found-that the variation in'measured losses is, infact, closely represented by the variationin C as derived from thegeneral theoretical equation given above, which would appear inconsequence to be valid, at least in the case ofa single stage turbine.

In explanation of the disparity of the-secondary losses estimated on apractical basis in a typical compressor stage and on a practical basisin a turbine stage, or on a theoretical basis in either case, it isconcluded that these losses are least in the case of the 'typical'intermediate stage of a multi-sta'ge compressor because in that caseonly is the blade stage followed immediately by a further relativelymoving blade stage. It is thought that the turbulence giving rise tosecondary loss is curtailed by that further blade-row instead ofextending some distance (in theory an infinite distance) downstream.From this conclusion one would expect the secondary losses in the finalblade row of a multi-stage compressor to be higher than in a typicalintermediate row; i. e. if in the latter' case the ratio of primary tosecondary losses is 6z'4,-inthe former case the ratio would be 624 3 or6:12, and consequently the ratio of total (primary plus secondary)losses of the final to any other row should be approximately6+12/6-l-j4, i. e. 1.8. This figure is effectively confirmed by the onlypublished measurements of losses through the various blade rows of'acompressor, where the correspondingratiois approximately 2. 4

-'It may, therefore, be assumed with confidence that while the primarylosses of any given row" of blades are substantially unaffected-byexternal influences, the secondary losses are considerably reduced by"the'mere presenceof a succeeding relatively"mov ing blade row, {thelatter actingas aloss inhibitor to the "said row -of blades'tAccordingly the invention is based on the proposition, arising out ofthat assumption, that the total losses in the said blade row when notsucceeded by another row will be greater than the combined total lossesof the said row and a succeeding (loss inhibiting) row provided thattheblades of the succeeding row have of themselves sufficiently lowcombined primary and secondary losses compared with the consequentreduction in secondary losses of the blades of the first mentioned row.

A quantitative illustration of this proposition is alforded byconsidering by way of example a particular working blade under certainoperating conditions with or without an inhibitor blade and using thecomparative loss relationships discussed in the foregoing. If it issupposed that the uninhibited secondary loss of the working blade wouldbe X then the inhibited secondary loss of the working blade would beX/3, and if it is supposed that the uninhibited secondary loss of theinhibitor blade would be y then the primary loss of the inhibitor bladewould be 6y/12 or y/Z. If, therefore, the total inhibitor blade lossesare to be smaller than the difference between inhibited and uninhibitedworking blade secondary losses then;

It would appear, therefore, that a net reduction in loss is possiblewherever a row of working blades, not being followed by a further row ofworking blades, can be followed immediately by a row of inhibitor bladeswhose losses are somewhat less than half as great in comparison withthose of the working blade. These inhibitor blades can moreover be ofsimple aerodynamic form with little or no camber so that their lossesare of a very low order, there being no other limitations affectingtheir design.

The invention offers the possibility of improving the efiiciency ofturbines in particular. It is well known that conventional turbines aregenerally of lower efficiency than compressors. The reason would appearat least partly to be that a typical compressor has several blade stagesof which only the last stage, usually a row of stator blades suffersuninhibited secondary losses which are of a relatively low order as theblades of the last row are not usually highly cambered; thus theuninhibited secondary losses are only a small proportion of the overalllosses. On the other hand, the typical turbine comprises fewer stages ofwhich the last stage, usually a rotor stage, has highly cambered blades,whose uninhibited secondary losses comprise a considerable proportion ofthe overall losses. The use of a row of inhibitor blades after the lastturbine rotor stage will, therefore, substantially reduce the overalllosses, giving a correspondingly greater efiiciency. The inhibitorblades are conveniently stationary and are so cambered, if at all, as totend to reduce any residual whirl component of velocity in the fluidleaving the last rotor stage.

The invention is considered to be of greatest importance in itsapplication to turbines having a single stage of rotor blading as arecommonly used in aircraft and other propulsion plants comprisingtypically a row of nozzle blades followed by a row of very heavilycambered rotor blades with, in some cases, a small number of strutsaffording support for a bearing or exhaust cone usually at some remotepoint downstream of the rotor blades. The secondary losses areparticularly heavy in such a turbine and are not sensibly inhibited bythe struts due to their remoteness or small number or both. Byproviding, in accordance with the invention, a row of inhibitor bladesof a numerical order comparable to or a large percentage of that of therotor blade row and immedia't'elfadjacent thereto, a significantreduction in chord leiigth of the blades of the first mentioned row.

in order that the, invention may be fully understood it will now.v bedescribed by way of example only in its embodiment in-a turbo-compressorjet propulsion plant with reference to the accompanying drawings inwhich:

Figure 1 represents a half cross section in a longi tudinal planethrough the jet propulsion plant;

Figures 2 and 3 represent fragmentary longitudinal and transversesections through the turbine blading of the plant of Figure l to anenlarged scale, Figure 3 being a section on the line III.III of Figure2;

Figure 4 represents a cross-section through the successive blade rows astaken along section line lVIV of Figure 2;

Figures 5(a) and (b) represent velocity triangles relating to the bladesections of Figure 4;

Figure 6 represents a cross-section through the successive blade rows ofan alternative form of blading to that of Figure 4;

Figures 7(a) and (b) represent velocity triangles relating to the bladesections of Figure 6.

In Figure 1 a turbine rotor 1 having a single row of blades 2 attachedthereto drives through a shaft 3 a com pressor rotor 4 carrying severalrows of blades 5 arranged in interdigital relationship with several rowsof stationary blades 6 attached to the compressor stator 7. Compressedair delivered at the annular compressor outlet 8 is supplied to anannular combustion system 9 wherein fuel is injected by the injector 10and the combustion gases are delivered through the flame tube 11 intothe turbine annulus 12. This annulus contains three rows ofcircumferentially evenly spaced blading disposed succesively andimmediately adjacent one another in the direction of fiow of thecombustion gases (shown enlarged in Figure 2) comprising firstly a rowof nozzle blades 13 attached to the turbine stator 14; secondly, therotor blades 2; and finally, in accordance with the invention, a row ofloss inhibiting blades 15 also attached to the turbine stator 14. Atransverse cross-section through this last row of blades 15 (at IlI-IIIin Figure 2) is shown in Figure 3 from which it will be seen that theblades are similarly closely spaced to those of the usual nozzle orrotor blade row. After traversing the blade rows 13, 2 and 15, the gasespass through the exhaust duct 16 to the jet orifice 17 where they issueas a high velocity propulsive jet.

The cross-sections of the blade rows 13, 2 and 15, at their mean radius(IV-IV in Figure 2) are shown in the fragmentary view of Figure 4, andthe corresponding velocity triangles for inlet to and outlet from therotor blades at the designed operating conditions are shown in Figures5(a) and 5 (b) respectively. The gases approach the nozzle blades 13 ina generally axial direction (indicated by the arrow A) and are deliveredtherefrom with a velocity V having a high circumferential component Vthe nozzle blades being highly cambered to effect this. Due to the factthat the circumferential velocity, U, of the rotor blades 2 is in thesame general direction as V the velocity of the fluid relative to therotor at inlets is reduced to V After traversing the rotor blades 2,which are necessarily highly cambered to derive sufficient energy fromthe fluid to drive the compressor, the gases leave with a relativevelocity V having in this case a circumferential component equal andopposite to the blade velocity U so that the absolute velocity of thefluid at outlet from the rotor, V,,, is purely axial in direction, beingequal in magnitude to the axial component of the absolute velocity atinlet to the rotor.

Consequently the gases enter the loss inhibiting blades 15 with a purelyaxial velocity, sothat these blades are accordingly of very simplesymmetrical streamlined section. It follows that the losses associatedwith these blades are small as compared with those of the highlycambered rotor blades which are inhibited by the presence of the former.Furthermore, although the rotor blades will normally be twistedthroughout their length to accommodate relative fluid velocityconditions at other radii ditfering from those at their mean radius,this need not be the case with the loss inhibiting blades since at allradii the fluid outlet velocity from the rotor blades can be arranged tobe entirely axial in direction. Accordingly the inhibiting blades 15 areconveniently of untwisted form and bi-convex in cross-section. By theterm untwisted, it is meant that the inlet and outlet blade angles areconstant along the blade lengm.

In the alternative form of blading of Figure 6, the nozzle blades 13 arethe same as in Figure 4 and the inlet velocity triangle to the rotor,Figure 7 (a), is also the same. In this case, however, the rotor blades2 are arranged so that the relative circumferential velocity componentof the fluid at their outlet, V in the outlet velocity triangle ofFigure 7 (b), is greater than the blade speed U so that the absolutefluid velocity at outlet from the rotor, V,,, has a smallcircumferential component, the axial component, 1,, being unchanged. Inthis case the loss inhibiting blades 15 are conveniently slightlycambered to eliminate the circumferential component of V so that thefluid velocity at exit from the loss inhibiting blades is entirely axialin direction. It will be apparent that a small degree of camber in theloss inhibiting blades 15 will not significantly detract from their lossinhibiting characteristics in association with the highly cambered rotorblades 2.

Owing to the very small camber, if any, of the loss inhibiting blades,they are not subjected to any large circumferential loads and the dragloads of the passing fluid are also necessarily small. Consequently, theattachment of these blades does not present any difiiculties. As shownin Figures 2 and 3 they are attached only at the outer end, beingpositioned and secured by lands 18 formed on their roots 19 which areengaged by complementary grooves in the associated abutting statorsections 14 and 20.

Although the exhaust cone 21 in Figure 1 is supported from the statorsection 19 by a small number of arms 22, this function could, of course,be performed by the loss inhibiting blade row.

What I claim is:

1. A turbine comprising a rotor partly defining a passage for the flowof working fluid, structure independent of said rotor also partlydefining said passage and with respect to which said rotor is rotatable,and a number of blade rows each having circumferentially evenly spacedelongate blades of streamlined profile in transverse section extendinglongitudinally across said flow passage which blade rows are arrangedsuccessively in the direction of flow of the working fluid, wherein thelast of said rows having blades shaped primarily for energy conversionand substantially cambered in section to derive considerable energy fromthe traversing fluid is on said rotor and is succeeded by .animmediately adjacent row of blades attached to said independentstructure and being in number a substantial proportion of the number ofthe blades of the preceding row and whose chord length is smaller thanthe chord length of the blades of the preceding row, said last mentionedblades which are attached to said independent structure being uncamberedin section.

2. A turbine according to claim 1, wherein the blades of said lastmentioned row attached to said independent structure have blade inletand outlet angles which are constant along the blade length.

3. A turbine comprising a rotor partly defining a passage for the flowof working fluid, stator structure also partly defining said passage,and a number of blade rows each having circumferentially evenly spacedelongate blades of streamlined profile in transverse section extendinglongitudinally across said flow passage which blade rows are arrangedsuccessively and immediately adjacent one to another in the direction offlow of the working fluid, of which the penultimate row traversed by theworking fluid is attached to said rotor and is the last of said rowshaving blades shaped primarily for energy conversion and substantiallycambered in section to derive considerable energy from the traversingfluid, wherein the blades of the last row traversed by the workingfluid, being in number a substantial proportion of the number of theblades of said penultimate row, are attached to said stator structureand have a substantially symmetrical biconvex section, are untwistedthroughout their length and having their chords lying in radial planescontaining the axis of said rotor.

References Cited in the file of this patent UNITED STATES PATENTS1,688,808 Gill Oct. 23, 1928 2,426,270 Howell Aug. 26, 1947 2,468,461Price Apr. 26, 1949 2,488,783 Stalker Nov. 22, 1949 2,613,869 AnxionnazOct. 14, 1952 2,648,492 Stalker Aug. 11, 1953 FOREIGN PATENTS 116,512Germany Dec. 31, 1900

